The disclosure relates generally to gas turbine engines, and, more specifically, to film cooling therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine (HPT), which powers the compressor, and in a low pressure turbine (LPT), which powers a fan in a turbofan aircraft engine application, or powers an external shaft for marine and industrial applications.
Engine efficiency increases with temperature of combustion gases. However, the combustion gases heat the various components along their flowpath, which in turn requires cooling thereof to achieve an acceptably long engine lifetime. Typically, the hot gas path components are cooled by bleeding air from the compressor. This cooling process reduces engine efficiency, as the bled air is not used in the combustion process.
Gas turbine engine cooling art is mature and includes numerous patents for various aspects of cooling circuits and features in the various hot gas path components. For example, the combustor includes radially outer and inner liners, which require cooling during operation. Turbine nozzles include hollow vanes supported between outer and inner bands, which also require cooling. Turbine rotor blades are hollow and typically include cooling circuits therein, with the blades being surrounded by turbine shrouds, which also require cooling. The hot combustion gases are discharged through an exhaust which may also be lined and suitably cooled.
In all of these exemplary gas turbine engine components, thin walls of high strength superalloy metals are typically used to reduce component weight and minimize the need for cooling thereof. Various cooling circuits and features are tailored for these individual components in their corresponding environments in the engine. For example, a series of internal cooling passages, or serpentines, may be formed in a hot gas path component. A cooling fluid, such as a relatively cool supply of compressed air, which may be supplied by the compressor of the turbine engine, may be provided to the serpentines from a plenum, and the cooling fluid may flow through the passages, exiting through one or more small holes formed on the wall surface, cooling the hot gas path component substrate and any associated coatings. However, this cooling strategy typically results in comparatively inefficient heat transfer and non-uniform component temperature profiles.
As indicated, in some instances, the supply of compressed air is released through small holes on the surface of the airfoils. Released in this manner, the supply of air forms a thin layer or film of relatively cool air at the surface of the airfoil, which both cools and insulates the part from the higher temperatures that surround it. This type of cooling is commonly referred to as “film cooling”. Film cooling involves a complex three dimensional flow. Interactions between a freestream and the cooling holes or jets, influence the overall film effectiveness. However, this type of film cooling comes at an expense. The release of the compressed air in this manner over the surface of the airfoil, lowers the aero-efficiency of the engine. In addition, the cooling fluid exiting the cooling holes into the freestream passage of high-temperature gas is easily separated from the wall surface, so that the efficiency of the film cooling is low. Current design technology is focused on shaped cooling holes, which use the geometry of the shaped hole to slow down and diffuse the film cooling leading to higher film effectiveness, yet can be costly to manufacture. As a result, there is an ongoing need for improved cooling strategies, including improved film cooling, for turbine airfoils.
It would therefore be desirable to provide a hot gas path component and method of forming cooling structures in the hot gas path component that provides for a more efficient and flexible cooling design that does not suffer from the above drawbacks.